Every metal part, no matter how well it is manufactured, contains some flaws as a result of the manufacturing process. The flaws are randomly distributed, some are on the surface and some are located internally. Flaws manifesting close to a stress concentrator (Kt) are subjected to increased stress and therefore have an increased probability of developing micro-cracks during service. These micro-cracks then subsequently join with each other to form a bigger crack. When the part is subjected to in-service loads, the crack begins to grow.
Metal fatigue has long been a problem for the aircraft industry. In the early 1950's, the first commercial jet airliner to reach production, the de Havilland DH 106 Comet, suffered several catastrophic accidents before the cause was attributed to metal fatigue. While a great deal has been learned since the 1950's about metal fatigue, and length-of-service standards have been established for parts replacement, even today the useful life of an aircraft part is not always well quantified. Indeed, in 2011, metal fatigue likely caused a five foot hole to rupture in the fuselage of a Southwest Airlines Boeing 737 resulting in an explosive decompression at 34,000 feet that necessitated an emergency landing.
The rate at which a micro-crack grows depends upon several factors that include load levels and the number of load cycles to which the material is subjected. Fatigue may be defined as the progressive and localized damage that occurs when a material is subjected to repeated load cycles. The inception of a crack is much debated in the literature and therefore there is no universal definition of crack initiation.
The United States Navy defines a crack to have initiated when the incipient crack reaches the length of 0.010 inch. The crack grows with applied load, eventually reaching a critical size at which point the part fails catastrophically. There are various non-destructive inspection techniques (NDITs) that allow the measurement of cracks, but only after reaching a certain size do cracks become detectable. FIG. 1 shows the general nature of crack growth, where the points on the graph: (ai, Ni), (ad, Nd), and (acrit Ncrit) define crack initiation (CI), threshold of detection, and critical points, where a is the crack length and N is the number of cycles it took for the crack to reach the length of a. Inspections are generally scheduled between Nd and Ncrit. For some aluminum alloys subjected to a fighter aircraft spectrum, approximately two-thirds of the total life is spent in the CI phase. The crack remains undetectable for some more time from Ni to Nd. The crack growth period from initiation in some high strength steel alloys is even shorter than aluminum, making it difficult to employ available NDITs. Thus there is a need to develop ways to detect cracks between Ni and Nd.
Crack initiation for a given stress is determined by laboratory testing of several dog-bone-shaped coupons (a coupon is a sample of a metal or metalwork) made from the metal or alloy in question. The coupons are subjected to a constant amplitude stress cycle with a minimum stress of 0.1 of the maximum stress for the metal to be tested (ratio (R) of 0.1). These tests are repeated for different stresses and different stress ratios to get a family of life curves. This method of determining CI life is referred to in the industry as “Method 1.”
Compact test specimens are typically used in studying crack growth. When such specimens are subjected to stress cycles with marker cycles introduced periodically, the crack leaves a pattern that can be studied under a microscope after the tests are completed. Such fractographic examinations can help chart the crack back to its origin. Using the Navy's definition, CI life can be determined. This method of determining CI life is referred to in the industry as “Method 2.” The CI lives determined by the above two methods may not match because of the lack of a physics-based definition of initiation. Also, there is some amount of uncertainty associated with CI life for a constant amplitude spectrum. Even more uncertainty is associated with CI life when the component is subjected to a variable stress spectrum due to in-service loads such as may be encountered by an F-18 while landing or maneuvering. Therefore, some amount of uncertainty is expected for any measurement technique, especially in the neighborhood of CI life and generally below the threshold of detectability of current NDIT.